Thermal elastic shock (TES) that spacecraft can experience is attributed to a disturbance torque created by the presence of a sunlit and a shadowed section of the orbit plane. The darkened region is formed as the Earth moves between the satellite and sun, thus eclipsing the satellite view of the sun. The eclipsed portion of the orbit is referred to as the umbra and the transition to and from this region is called the penumbra. Thermal elastic shock is a result of a rapidly changing temperature difference between the hot (sun pointing) and cold (anti-sun pointing) surfaces of a large flexible appendage extending away from the spacecraft. The temperature offset across the two surfaces quickly decreases at umbra entrance when the solar heat flux is turned off in a stepwise fashion, thus subjecting the flexible structure to a rapid cooling effect. In a similar respect, the thermal gradient quickly increases as the appendage experiences rapid heating during umbra exit when the solar flux turns on in a stepwise manner. The rapid change of the thermal gradient causes the hot surface of the structure to bend due to either thermal compression (sunset) or thermal expansion (sunrise). The thermally induced bending motion of the flexible member generates a disturbance torque which is then transferred back onto the space vehicle core-body.
To minimize or eliminate the TES disturbance torque, feasible solutions, other than the solution offered by this invention, do exist. One obvious solution is to alter the material properties of the disturbing member so that the thermal gradient at umbra entrance and exit is not so dramatic. However, this solution must be pursued during the design phase of a spacecraft before the hardware development begins. If the hardware phase is already underway, redesigning the structure will be quite costly. Furthermore, prior to the research leading to this invention a detailed TES disturbance model was unavailable to determine the effect of the TES disturbance on the spacecraft attitude pointing performance. Thus an accurate upper bound of the allowable thermal gradient to maintain pointing accuracy was also unavailable.
Another solution is to impose structural constraints on the appendage to increase rigidity and structural damping. As in the case of the first alternative solution, this solution also requires additional analysis to determine the required stiffness to minimize pointing errors introduced by the TES disturbance. Although this solution may not be as costly as the first alternative approach, some additional costs would be involved. Again, the use of the TES disturbance model is needed to produce estimates of the vehicle attitude pointing performance.
A final approach would be to use feed forward compensation techniques in the vehicle attitude control system to minimize the effect of the TES disturbance. Although this solution seems to be the easiest and cheapest to employ, it can be quite deceiving. This approach would require a software change in the vehicle on-board computer and some resource to obtain thermal gradient measurements. The TES disturbance equations would be an integral part of the software needed to predict the magnitude of the TES disturbance. However, current reaction wheel assemblies (RWA's) reach their maximum torque authority responding to the TES disturbance as evidenced and predicted for some spacecraft (i.e. TOPEX, LANDSAT). Thus the feed forward torque prediciton method may provide little or no compensation. An actuator capable of generating larger torques would be necessary for the feed forward compensation technique to be valid. Unfortunately, current RWA designs do not produce enough torque and new RWA's would be costly to develop. Large torque actuators, such as a control moment gyroscope (CMG), could provide the necessary torque to eliminate the TES disturbance, but CMG's are bulky, heavy and require a large power input. Also, for small satellites, CMG's are not a practical solution.
All of the preceeding described solutions would only be valid for the individual spacecraft under consideration. Furthermore, these solutions would require the use of the previously mentioned TES torque equations. The spacecraft thermal disturbance control system invention overcomes these previous problems. With this invention it is possible to effectively counter the effects of thermal elastic shock without changing the materials used in the manufacture of portions of the spacecraft. Moreover, this invention does not require the redesign of the spacecraft. In fact, it is possible to retrofit many spacecraft to use this invention without any extensive redesign or increases in cost. The localized active control system solution of this invention on the other hand, would still require the use of the TES disturbance model, but would not be limited to a specific vehicle.